The present invention relates to an airfoil for use in the turbine section of a gas turbine, such as in a stationary vane or rotating blade airfoil. More specifically, the present invention relates to a cooling air regulating seal for use in a gas turbine airfoil.
A gas turbine employs a plurality of stationary vanes that are circumferentially arranged in rows in a turbine section. Since such vanes are exposed to the hot gas discharging from the combustion section, cooling of these vanes is of utmost importance. Typically, cooling is accomplished by flowing cooling air through cavities, such as fore and aft cavities, formed inside the vane airfoil. A tubular insert is typically disposed in each of these cavities so that passages surrounding the inserts are formed between the inserts and the walls of the airfoil. The inserts have a number of holes distributed around their periphery that distribute the cooling air around these passages.
A major portion of the cooling air flowing through the aft cavity is typically discharged through a cooling air passage in the trailing edge of the airfoil. Although baffles extending between the insert and the airfoil wall have sometimes been used to control the direction of the flow of the cooling air around the passages that surround the insert, the cooling air in those passages was still allowed to flow freely to the passage in the trailing edge for discharge from the airfoil--that is, the flow of cooling air from the passages that surround the insert to the trailing edge passage was not positively regulated. Consequently, the pressure inside the passages surrounding the inserts was set, by and large, by the flow capacity of the trailing edge passage.
Another portion of the cooling air flowing through the aft cavity is typically directed to film cooling air holes in the concave wall that forms the pressure surface of the airfoil near the trailing edge. These film cooling holes direct the cooling air over the vane airfoil pressure surface so as to provide a degree of film cooling near the trailing edge.
The flow rate of the cooling air through the film cooling holes is a function of the pressure differential between the cooling air flowing through the passages surrounding the insert that supply the film cooling holes and the hot compressed gas flowing over the pressure surface of the airfoil. In some modern high performance gas turbines, the gas loading on the vane airfoil is relatively high, thereby reducing this pressure differential. Unfortunately, the essentially unregulated flow of cooling air from the passages surrounding the aft cavity insert to the discharge passage in the trailing edge, discussed above, may reduce the pressure of the cooling air in the cavity surrounding the insert to the point where the pressure differential between it and the hot compressed gas becomes too low to provide sufficient cooling air through the film cooling holes. This situation can result in overheating of the airfoil pressure surface in the vicinity of the trailing edge.
One potential solution to this problem is to dramatically increase the cooling air supplied to the airfoil cavity, thereby increasing the pressure of the cooling air flowing through the passages surrounding the insert. However, such a large increase in cooling air flow is undesirable. Although such cooling air eventually enters the hot gas flowing through the turbine section, little useful work is obtained from the cooling air, since it was not subject to heat up in the combustion section. Thus, to achieve high efficiency, it is crucial that the use of cooling air be kept to a minimum.
It is therefore desirable to provide a scheme for increasing the pressure of the cooling air in the passages surrounding the insert without increasing the flow of cooling air through the cavity.